The use of composite materials in the manufacture of aircraft and other lightweight structures has increased steadily since the introduction of such materials. Composite materials have a high strength-to-weight ratio and high stiffness, making them attractive for use in lightweight structures. Two drawbacks to using composite materials are their high fabrication costs and low damage tolerance. It is difficult to produce composite parts for the same cost and having the same degree of damage tolerance as comparable metal parts. This cost differential is especially notable in large scale composite parts having complex contours.
Some contributors to the cost of composite structures are the amount of manual labor required during fabrication and the cost of the complex tools generally used to form composite parts having complex geometries. Complex parts having complex geometries, for example, aircraft wing boxes, are generally formed by manually laying up individual layers of composite material on a contoured mandrel to form the exterior skin of the wing box. The spars and ribs of the wing box are generally formed separately by manually laying up individual layers of composite material on a series of different tools in order to form individual I-beam or sine wave spars and ribs. The sine wave spars or ribs are generally vacuum bagged and precured as separate components within an autoclave under elevated temperatures and pressures. In one approach, after curing, the sine wave spars and ribs are placed on the laid up outer skin of the wing box and subsequently adhesively bonded to the skin in an autoclave operation in which the skin, spars and ribs are placed in a vacuum bag and cured at an elevated temperature and pressure. In order to insure good structural integrity between the spars, ribs and skin, "chicken fasteners" are generally used to reinforce the spar/rib-skin interface.
In order to decrease fabrication costs and increase structural integrity, it has been found advantageous to join the spars and ribs to the skin prior to curing the individual composite substructures. The skin and joined ribs and spars are then vacuum bagged with appropriate tooling and co-cured together in a single autoclave operation. This method of fabrication eliminates the separate curing cycles required to form individual composite subassemblies and subsequently join them to form the wing box structure. However, currently it is difficult to properly design and fabricate appropriate tools to jointly cure complex structures while maintaining structural tolerances and minimizing or eliminating areas of resin porosity or resin richness.
One problem that is encountered in forming spars and ribs to a skin in one autoclave operation is that the complex geometries formed by the skin and ribs are difficult to support by tooling. The spars typically are of an I-beam construction, and the webs of the spars extend substantially perpendicular to the skin. One end cap of the I-beam spar extends against the face of the skin. The other end cap of the spar extends parallel to the skin, but spaced apart from the skin. Ribs extend perpendicular to both the spars and the skin. Five-sided cavities are formed by a pair of adjacent spars, a pair of adjacent ribs, and the skin. Adjacent outer end caps on adjacent spars create a partially closed opening for the sixth side of the cavities into which tooling must be inserted to support spars during the autoclave process.
Inserting tooling into a cavity with such a partially closed opening can be difficult, but removing the tooling from the partially closed opening after curing of the spars and the skin is nearly impossible. One prior art manner of dealing with this problem developed by Lockheed Martin involved upper and lower cure fixtures for the skins and spars. The composite material for the skins were laid up against the lower skin cure fixture. The composite material for the spars and ribs were laid up over rubber sub-bladders having the shape of the to-be-formed cavity, and then were arranged over the skins. An upper spar cure fixture was then aligned with the top of each of the upper end caps of the spars. The upper spar cure fixture included a groove along which the upper end cap was seated. The upper spar cure fixture also included vacuum attachments and a seal that received a bladder bag. The bladder bag was shaped and configured so that it could extend into the cavity formed between adjacent spars and ribs and the skin. The soft structure of bladder bag and subbladder facilitated easy tool removal after curing. However, the inner tooling and upper spar cure fixture were expensive to produce and cumbersome to use. Moreover, close inner part tolerances were difficult to produce because the soft tooling was difficult to maintain in the proper final shape of the cavity. There is a need for a more efficient manner of providing this inner tooling.
Another problem associated with past tooling designs is the inability of the tooling to place appropriate consolidation pressures on all of the regions of a complex part during curing, thus leading to areas of resin richness or resin porosity. Generally, past tooling concepts have been formed of various materials such as Invar 42 or composite materials that have similar coefficients of thermal expansion as the composite material from which the part is being formed. Matching the coefficients of thermal expansion between the tooling materials and the composite material from which the part is formed is thought to reduce problems associated with thermal expansion mismatches during part curing.
Prior tooling concepts, even those using tools formed from materials having similar coefficients of thermal expansion as the composite materials, often fail to achieve complete part consolidation. Consolidation problems are complicated in composite parts where not only must complete consolidation be achieved, but also part dimensional tolerances must be carefully controlled. Part tolerances are increasingly important in parts having both exterior mold line surfaces and interior ribs, spars or flanges whose geometry and locations must be carefully controlled.
Generally, prior tooling concepts for use in complex composite parts have relied upon matched metal tooling. However, there is even small amounts of misalignment in the location of any of the surfaces of the matched metal tooling, the result is often unacceptable parts having areas of resin richness or resin porosity. Alternatively, complex metal tools incorporating indexing substructures have been used. As with matched metal tooling, such tooling often fails to produce high quality composite parts, again due to the inability of matched metal tooling to compensate for even slight errors in position. In addition, the complexity of most such prior tooling concepts has resulted in expensive tools that contribute to the overall costs of the composite parts being produced.
One of the contributions to resin richness and resin porosity in cured composite parts is the large difference between the uncured and cured dimensions of the composite part. Because composite parts are formed of many layers of composite material that are joined together and then, consolidated and cured, the laid up but uncured composite workpiece has a much greater volume than the fully compacted and cured composite workpiece. This changing volume and thus part thickness during curing is complicated by the fact that the magnitude of compaction-consolidation of the composite workpiece differs at different areas within the workpiece due to the varying number of layers of composite material used to form different areas of the workpiece.
Prior tooling concepts are configured to account for the thermal expansion of the composite tools and the composite workpiece during curing. However, such tooling concepts fail to account for the varying magnitudes of compaction/consolidation required at different areas in the composite workpiece. Thus, prior tooling has been used very successfully to form composite parts having approximately constant thicknesses, but has not been used as successfully to form composite parts having varying thicknesses and close tolerances on the surfaces of the part.
As can be seen from the above discussion, there remains a need in the industry for improved composite fabrication methods and apparatus. The present invention is directed toward composite fabrication methods and apparatus that fulfill part of this need.